This invention relates to tilt rotor helicopters and specifically to helicopters having variable speed tilt rotors for achieving substantial increases in endurance, range, altitude and speed and reductions in noise levels and fuel consumption.
The efficiency of an aircraft, whether fixed wing or rotorcraft, as expressed by the fuel consumption required to achieve a specific performance as for example, cruise, climb, or maximum speed, is directly proportional to the power required to achieve such performance. The power required is inversely proportional to the ratio of the aircraft lift to the drag (L/D). In order to increase an aircraft efficiency designers strive to increase the lift to drag ratio by minimizing the aircraft drag at lift levels required to counter the aircraft weight and to allow for aircraft maneuvering.
The lift and drag of an aircraft are determined by the following formulas, respectively:
L=xc2xdxcfx81V2SCLxe2x80x83xe2x80x83(1) 
D=xc2xdxcfx81V2SCDxe2x80x83xe2x80x83(2) 
Where xcfx81 is the air density, V is the air velocity (airspeed), S is the reference area of the lifting surface (wing or rotor blade), CL and CD are non-dimensional lift and drag coefficients. The lift to drag ratio L/D is equal to coefficient of lift to the coefficient of drag ratio, CL/CD. Thus, the ratio of the coefficient of lift to the coefficient of drag, CL/CD, has a direct effect on performance. The CL/CD is a function of CL as can be seen by the CL V. CL/CD graph depicted in FIG. 1 for a typical airfoil. For best cruise efficiency, the coefficient of lift of the lifting airfoil should be maintained at levels of maximum CL/CD.
In a helicopter the lift and drag of the rotor blades conform to the same lift formula L=xc2xdxcfx81V2SCL where V is the local airspeed on the blade which, in a hovering helicopter is a result of the blade angular velocity in revolutions per minute (RPM). For convenience, xe2x80x9cRPMxe2x80x9d as used herein refers to rotor angular velocity. Moreover, the term xe2x80x9chelicopterxe2x80x9d as used herein encompasses all types of rotorcraft.
In a hovering helicopter, the speed of the rotor blade increases radially outward. At any given radial distance from the rotor center, the speed of the blade is given by the equation:                     vr        =                              2            ⁢            π            ⁢                          xe2x80x83                        ⁢                          r              ⁡                              (                RPM                )                                              60                                    (        3        )            
where, vr is the rotational speed and r is the radial distance measured from the rotor center.
A helicopter in a substantial forward speed (e.g., 100-200 mph) experiences problems of control, vibration and limitations in performance resulting from the asymmetry in the speeds of the advancing and retreating blades. When traveling in a forward direction 8, the advancing blade 10 has a speed equal the rotational speed of the blade plus the forward speed of the helicopter, whereas the retreating blade 12 has a speed equal the rotational speed of the blade minus the forward speed of the helicopter. The speeds along the length of the blades when traveling forward are shown in FIG. 2. As a result, the advancing blade has more lift than the retreating blade. To avoid helicopter roll over due the airspeed asymmetry, the lift on the retreating blade has to be increased while the speed on the advancing blade has to be decreased. Because, lift is inversely proportional to the velocity (i.e., speed) of the blade squared (V2) a substantial increase in the coefficient of lift (CL) of the retreating blade is required. The available lift coefficient for a given blade is limited as shown FIG. 1. Consequently, the asymmetry in speeds between the advancing and retreating blades has to be limited thereby limiting the forward speed of the helicopter.
Increasing the RPM of the rotor reduces the relative asymmetry of the airspeed distribution, thus reducing the effects of forward speed on roll control limits. But such RPM increase is constrained by the maximum allowable rotor tip speed. The maximum allowable tip speed is typically lower than the speed of sound (i.e., Mach 1) so as to avoid the substantial increases in drag, vibration and noise encountered when the tip speed approaches Mach 1.
Current helicopter rotors turn at a constant RPM throughout the flight because of the complex and severe rotor dynamics problems. Generally, helicopter designers are content if they succeed in the development of a single speed rotor, which can go from zero to design RPM when not loaded on the ground during start and stop without encountering vibration loads which overstress the helicopter and rotor structure. When the blades of a conventional rotor are producing lift, a significant change of the rotor blade RPM from the design RPM may yield catastrophic results.
Conventional helicopter rotors are designed to achieve blade flap, lag and torsional natural oscillation frequencies, at the operating RPM, which are adequately separated from the rotor excitation frequencies occurring at the rates of 1 per revolution, 2 per revolution, 3 per revolution and so forth. For example, for a rotor operating at 360 RPM, the frequency corresponding to the occurrence of a rotor excitation frequency of 1 per revolution is 6 Hz (360 RPM is 6 cycles per second), 2 per revolution is 12 Hz, and so forth. As the rotor RPM is changed so are the excitation frequencies. For convenience, the frequencies which give rise to these excitation frequencies are referred to herein by the excitation frequency occurrence rates. For example a frequency that gives rise to an excitation frequency that occurs at a rate of 2 per revolution is referred to herein as the xe2x80x9c2 per revolutionxe2x80x9d frequency. For good dynamic behavior, considering both blade loads and helicopter vibration, conventional rotors with any number of blades are designed to avoid the frequencies of 1 per revolution, 2 per revolution, 3 per revolution and so forth. Conventional rotor blades are designed to operate at 100% of design RPM with the fundamental flap mode at a frequency above the 1 per revolution frequency, the fundamental lag mode usually below the 1 per revolution frequency and sometimes between the 1 per revolution and the 2 per revolution frequencies, and the blade dynamics tuned so that higher flap, lag torsion modes avoid the 1, 2, 3, 4, . . . n per revolution frequencies. The conventional blade design modes (i.e., modal frequencies) must be kept separated from the 1, 2, 3, 4, . . . n per revolution frequencies to avoid the generation of vibration loads which may be catastrophic. As a minimum, such vibration loads will make the helicopter unacceptable for the pilot and passengers and detrimental to the reliability of its mechanisms and equipment. To avoid such vibration loads, the rotor angular velocity is limited to a narrow range around 100% of design RPM, except for start-up and shut-down at low or no rotor load and low wind speed.
The RPM of helicopter rotors is normally set for a maximum forward speed at a maximum weight at a certain critical altitude. The RPM of the rotor is such that at maximum forward speed, the tip of the advancing blade is traveling at speeds near but below Mach 1, to avoid the substantial increases in drag, vibration and noise encountered at speeds approaching Mach 1. At any other flight conditions, the rotor RPM and thus, the power required to turn the rotor are substantially higher than that required for efficient operation.
Some research helicopters such as the Lockheed XH-51A compound helicopter have experimented with rotor RPM reduction at certain flight conditions by incorporating a wing for producing most of the required lift and a jet or a propeller driving engine for producing the required forward thrust. The use of the wings and engine relieve the rotor of its duty to produce lift and thrust, thus allowing the unloaded rotor to operate at reduced RPM. In this regard, a helicopter can fly at higher speeds before the tip of the advancing blade approaches the speed of sound and encounters the increased levels of vibration and noise as well as drag.
Other attempts have been made in improving helicopter maximum forward speeds and/or reducing noise at maximum speed by using 2-speed gearboxes. These gearboxes allow the rotor to rotate at two RPM values while maintaining a constant engine RPM. The rotor is set to rotate at a lower RPM when at high forward speed so as to reduce the rotor tip speed. In all other conditions, the rotor is set to rotate at the higher RPM. However, these attempts do not substantially improve the efficiency of the helicopter by reducing fuel consumption.
Another helicopter uses 10% reduction in rotor RPM during takeoff and landing in order to conform to very strict noise limitations. Because of this reduction in rotor RPM, the helicopter performance is compromised during take-off and landing.
While these aforementioned endeavors attempted to increase maximum speed and reduce noise during take-off and landing, neither attempted to improve the efficiency of the helicopter. Neither attempted to reduce the fuel consumed and power required for a given performance or attempted to increase a helicopter performance without increasing the fuel consumed and the power required. As such there is a need for a helicopter rotor system which will improve helicopter range, altitude and speed performance while reducing fuel consumption and noise levels.
Tilt rotor type rotorcraft incorporate wings which produce lift in forward flight, and, at forward speeds which is adequate to support the weight of the rotorcraft. The rotors (usually 2 or 4) are xe2x80x9ctiltedxe2x80x9d from a first position where their axis of rotation is vertical and where the rotors act as a regular helicopter rotors to a second position where their axis of the rotation is relatively horizontal and the rotors act as propellers producing forward thrust. A tilt rotor type rotorcraft converts from helicopter mode to airplane mode (wing borne with propellers) after vertical take-off and converts back to helicopter mode for hover or vertical landing.
The best-known tilt rotor rotorcraft is the V-22 Osprey. The V-22 Osprey uses 2-speed rotors, a 412 RPM for helicopter mode and for conversion to airplane mode and 333 RPM when the rotors are locked in propeller mode for forward flight.
The experimental predecessor to the V-22, the XV-15, attempted the same type of 2-speed rotor but was not successful in achieving such 2-speed because of rotor dynamics, which caused high vibration and loads at RPM other than 100%.
The constant RPM or the close-ratio (100% and 81%) RPMs of the current tilt rotors result in excessive RPM in forward flight and excessive blade xe2x80x9ctwistxe2x80x9d (variance of blade angle at the tip of the blade vs. the angle at the root of the blade) for hover flight. Both of these limit the performance and efficiency of the current tilt rotors. As such, there is a need for a tilt rotor system which will improve the tilt rotor type rotorcraft maximum hover weight, cruise range, altitude and speed performance while reducing fuel consumption and noise levels.
The present invention provides a variable speed tilt rotor and a method for using the same for improving tilt rotor rotorcraft performance and efficiency while reducing fuel consumption. Both in helicopter mode and in airplane mode, the RPM of the rotor system of the present invention can be varied to multiple and even infinite settings depending on the rotorcraft flight conditions to maintain a blade loading for optimum performance and fuel efficiency. In helicopter mode, the present invention allows for reduced rotor RPM at reduced forward speeds achieving an increase in rotor blade lift coefficient at the lower forward speeds and higher blade lift to drag ratio and thus, higher aerodynamic efficiency, lower required power, fuel consumption and noise level. By decreasing the RPM of the rotor, the power required to drive the rotor at the decreased RPM is also decreased. The adjustment to rotor RPM and power can be accomplished manually or automatically as for example by computer.
In airplane mode, the present invention allows a major reduction in RPM and increase in propeller efficiency, rotorcraft endurance, range, altitude and speed. The current invention allows dramatic reduction of the weight of the rotor blades, which makes possible large rotor diameters unachievable with current rotors. The combination of very low RPM in airplane mode and the low blade mass virtually eliminates whirl mode flutter and therefore allows for higher efficiency wings (larger span and narrower chord), thereby further increasing endurance, range, altitude and reducing fuel consumption and noise levels.
In order to be able to operate over a wide RPM range, the tilt rotor system of the present invention is designed to be able to operate close to or on rotor excitation frequencies. To achieve such unique capability, the rotor blades are designed to be stiff and extremely lightweight. The blades should be substantially lighter than conventional rotor blades.
In an exemplary embodiment, the tilt rotor blades flap, lag and torsion stiffness as well as the blade weight per unit length are continuously decreasing from the blade root to the blade tip. Applicant discovered that to achieve efficient operation at a wide range of tilt rotor angular velocities, the blades of the present invention preferably have a flap stiffness and a blade weight as follows:
Flap Stiffness: EIflapxe2x89xa7200 D4 at 30% of rotor radius measured from the center of rotor rotation
Total Blade Weight: Wxe2x89xa60.0025D3 
where D is the rotor diameter and is measured in feet, W is the total weight of each blade in pounds, and EI is in lbs-in2.
In the exemplary embodiment, the tilt rotor is not gimbaled, like conventional tilt rotors, but the blades are rigidly attached to the rotor hub in the flap and lag directions (without requiring hinges, elastometers or flexbeams) and the hub is rigidly attached to the mast. A blade bearing system is provided to change the blade angle of incidence around the feathering axis. This non-gimbaled hingeless rotor provides for substantial increase in rotor control moments using cyclic controls. Cyclic control is the control of the blades on the one side of the rotor disc to provide higher lift/thrust than those on the other side by changing the angle of incidence of the blades around the feathering axis. The OSTR provides for higher rotor control moments in the pitch and roll directions in helicopter mode and in the pitch and yaw directions in airplane mode and thus for substantial increase in rotorcraft maneuver and for flight with wider range of the rotorcraft center of gravity. The powerful control capability in pitch and yaw in airplane and conversion modes makes the tail unit, both vertical fin and horizontal tail, completely unnecessary or they can be drastically reduced in size with a resulting substantial reduction in rotorcraft weight and drag.
Another advantage of high stiffness OSTR rotor is that it enables the additional use of Individual Blade Pitch Control (IBPC) to provide gust and rotor load alleviation.